(1) Field of the Invention
The present invention relates to a rotorcraft tail rotor. The term rotorcraft is used in this application to describe any type of rotary wing aircraft, such as in particular helicopters.
(2) Description of Related Art
In current usage, such a tail rotor is also called an “anti-torque” rotor since it makes it possible to exert a torque for opposing the rotary movement in yaw produced by a main rotor, which provides the rotorcraft with lift and propulsion.
Thus, the tail rotor of a rotorcraft generally has a substantially horizontal axis of rotation and it may either be integrated in the fuselage of the rotorcraft, or else it may be external thereto. When the rotor is integrated in the fuselage, it is then referred to by the terms ducted tail rotor or Fenestron®.
In addition, the tail rotor may comprise at least two blades arranged symmetrically about the axis of rotation of the rotor. Each blade is movable in pivoting about an axis referred to as the “pitch” axis in order to vary the angle of incidence of the blades relative to the surrounding air and thus vary the thrust of the rotor. Such variation can in particular be carried out when the pilot of the aircraft wishes to modify the yaw angle described by the aircraft, or more simply to increase the power of the main rotor, with a consequent need to increase the value of the yaw torque.
Furthermore, in order to vary the pitch angle, the pilot actuates pedals of a rudder bar making it possible to actuate the rod of a tail rotor servo-control that is hydraulically connected to a control plate. Such a control plate can be moved in translation in a direction parallel to the axis of rotation of the rotor. Moving the pedals, from one abutment to another, varies the pitch of the blades of the antitorque rotor, by means of each rod connecting said plate to each blade, through an angle that may for example lie in the range −8° and +23° about a flat position of 0°.
In flight, as soon as the thrust of the rotor is no longer zero, i.e. when the pitch angle is different from 0°, torque for returning the blades to a flat position is exerted on the blades and tends to return the pitch angle to 0°. When a rotor is in rotation, the centrifugal forces acting on each blade element cause any pitch variation to be opposed, and the blades to be returned to the plane of rotation. In order to calculate the resulting torque M being exerted on the blades, reference may be made to the known literature on the subject, and for example, consideration may be given to a mass element dm of the blade, situated at a distance r′ from the center of the rotor. Then, the centrifugal force element acting on that mass element has the value:dF=Ω2·r′·dm where Ω is the angular velocity of the rotor.
This force element can then be resolved into a component dFZ parallel to the pitch axis and a component dF1 perpendicular to the pitch axis. The component dF1 then has the value:dF1=dF·sin φ=Ω2·r′ sin φ·dm=Ω2·x·dm where x is the distance of the blade element under consideration from the pitch axis.
The component dF1 is then situated at a distance y from the plane for which the general pitch angle θ is zero, thus giving rise to a torque element dM tending to return the blade to a flat position, i.e. in a plane xOz.
However, if it is accepted that the center of gravity G of the section of the blade is situated on the pitch axis, the value of that torque element is given by the expression:dM=Ω2·x·y·dm and the resulting torque M is written:
  M  =            Ω      2        ·                  ∫                              -            C                    ⁢                                          ⁢          1                          C          ⁢                                          ⁢          2                    ⁢              x        ·        y        ·                  ⅆ          m                    with C1 and C2 corresponding respectively to the orthogonal projections onto an axis Ox of a leading edge of the blade section at the point G and of a trailing edge of blade section at the point G.
With hydraulic power assistance, and more particularly by means of the servo-control, the “return-to-flat” effect of the rotor can be countered so as to eliminate the reversibility of the command. However, in the event of failure of such hydraulic assistance, as may for example result from a leak in the hydraulic system or from a failure of a hydraulic pump, the force necessary to operate the pedals suddenly becomes very considerable.
For that reason, an additional hydraulic safety device, referred to as a yaw “force compensator” is installed, along with the servo-control, which makes it possible to offload the major portion of the aerodynamic force feedback generated by the tail rotor. Such a force compensator makes it possible to generate a force opposing the return-to-flat force being exerted on the blades. The force compensator thus comprises a hydraulic pressure accumulator that is independent from the main hydraulic assistance system and a control lever for multiplying the movement stroke of a piston of an actuator that is connected hydraulically to the pressure accumulator in order to create a “spring having a negative slope”.
However, such a force compensator also incorporates numerous hydraulic members that can also be subjected to damage. A leakage in the force compensator would then prevent production of the force necessary for opposing the return-to-flat force.
That could therefore result in blocking of the system for controlling pitch variation of the tail rotor. Specifically, in the event of simultaneous malfunctions of the hydraulic assistance and of the force compensator, the forces required to enable the pilot to modify the pitch of the blades of the tail rotor are then too great and they prevent the pilot from varying the pitch of the blades of the tail rotor by using the pedals.
Thus, a first object of the invention is to limit possible causes of control over the pitch angle of a tail rotor becoming blocked, and thus to improve rotorcraft safety.
In addition, with increasingly powerful aircraft engines, it has become necessary to increase the surface area of the blades of the tail rotor, e.g. by adding “tabs” to the trailing edge of each blade. This consists in adding elements to the trailing edge, which elements extend a few centimeters in the chord direction and occupy a longitudinal fraction of the span. This increase in the surface area of the blades thus generates a considerable increase in the static and dynamic force transmitted by the blades to a rotor head.
In order to reduce these forces, the root of each blade is fitted with two compensation weights emerging symmetrically substantially perpendicularly to a main inertia axis of each blade, or more simply to the longitudinal mid-plane defined by each blade. Those compensation weights serve to create a moment opposing the return-to-flat moment of the blades and thus to reduce the forces required to control the pitch angle. Each blade element is thus stabilized, regardless of the pitch angle of the rotor.
Such compensation weights are more generally referred to as “Chinese weights”. They thus co-operate with each blade to form a rigid single-piece unit, and in particular they are described by the Applicant in document FR 2 719 554. Indeed, that document describes the compensation weights as being stationary elements forming projections on both sides of a longitudinal mid-plane of the blade.
However, although such Chinese weights make it possible to limit static force, they do not make it possible to reduce the dynamic force transmitted by the blades to a rotor head. Such constraints may then lead to reducing the lifetime of the revolute joint between each blade and a hub body. Such a revolute joint is indeed formed by elements constituting the pitch hinge, formed by laminated bearings. Those bearings are constituted by a combination of metal and elastomer, and they are generally cylindrical or even conical in shape. Furthermore, they are the site of considerable mechanical stress during rotation of the tail rotor, and more particularly during stages of varying the pitch angle.
That thus results in maintenance intervals for the aircraft being shortened, in particular concerning replacement of the wear members that take up these forces. However, shortening maintenance intervals increases the cost of operating those aircraft, and commercially that is to be avoided.
In addition to Chinese weights that are stationary relative to the roots of the blades of a tail rotor, it is also known to fit a main rotor of a rotorcraft with oscillating pendulums or weights that are movable in pivoting relative to the blade roots. Such arrangements are described in particular in documents FR 2 530 216, FR 2 435 391, and FR 2 959 484 but they do not make it possible to guarantee good reduction of the static and dynamic force generated by the rotation of a tail rotor.
Indeed, the projection described in document FR 2 530 216 is formed by the casing of the shaft 6 and is thus hollow. Such a casing thus does not act as a Chinese weight in the same way as a solid projection. Furthermore, the direction of the revolute joint between the pendulum and the casing is perpendicular to the direction in which the casing emerges relative to the main inertia axis of the blade element. Such an arrangement is therefore not suitable for reducing the static and dynamic force generated by the rotation of a tail rotor.
Document FR 2 435 391 describes a main rotor of a rotorcraft provided with weights oscillating relative to a blade element. However, strictly speaking, there is no projection emerging perpendicularly to a main inertia axis of a blade element. The oscillating weights are thus directly positioned on either side of the blade element, without being spaced apart from a main inertia axis. Such an arrangement is therefore not suitable for reducing the static and dynamic forces generated by the rotation of a tail rotor in simple and optimum manner.
Document EP 0 058 117 describes a suspension for a main gearbox of a helicopter with oscillating weights connected to the fuselage by deformable portions, but it is not transposable to a rotorcraft tail rotor for the purpose of reducing the static and dynamic forces generated by the tail rotor rotating.
Furthermore, in another alternative for limiting stress in the laminated bearings, it is also possible to limit engine power, and thus aircraft speed, which is also prejudicial from a commercial point of view. However, such a solution is merely palliative and does not under any circumstances enable the problem to be resolved at its source.
A second object of the present invention is thus to provide a rotor that enables the above-mentioned limitations to be overcome, and in particular that significantly reduces the static and dynamic forces generated by the tail rotor rotating. Thus, the structural design of the tail rotor of the invention makes it possible to limit, or even to eliminate, the mechanical stress transmitted to the laminated bearings, to the pitch control rods, and to the entire drive linkage, and does so while using the engine(s) at maximum power.
In addition, as described in document EP 0 773 881, tail rotors are also known in which a gyroscopic mechanism makes it possible to vary thrust automatically. It is thus possible to stabilize the yaw torque of the helicopter in flight.
However, such a solution is mechanically complex to implement and thus generates considerable manufacturing and/or adaptation costs in comparison with current solutions involving Chinese weights, hydraulic assistance by servo-control, and/or a force compensator.